In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine, upon which the hot combustion gases impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at metal temperatures of up to about 1900-2100.degree. F.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes to their current levels. For example, the composition and processing of the base materials themselves have been improved.
Physical cooling techniques may also be used. In one technique, internal cooling passages through the interior of the turbine airfoil are present. Air is forced through the cooling passages and out openings at the external surface of the airfoil, removing heat from the interior of the airfoil and, in some cases, providing a boundary layer of cooler air at the surface of the airfoil. To attain maximum cooling efficiency, the cooling passages are placed as closely to the external surface of the airfoil as is consistent with maintaining the required mechanical properties of the airfoil, to as little as about 0.020 inch in some cases.
In another approach, a protective layer or a ceramic/metal thermal barrier coating (TBC) system is applied to the airfoil, which acts as a substrate. The protective layer with no overlying ceramic layer (in which case the protective layer is termed an "environmental coating") is useful in intermediate-temperature applications. The currently known protective layers include diffusion aluminides and overlays. A ceramic thermal barrier coating layer may be applied overlying the protective layer on the external airfoil surface, to form a thermal barrier coating system (in which case the protective layer is termed a "bond coat"). The thermal barrier coating system is useful in higher-temperature applications. The ceramic thermal barrier coating insulates the component from the combustion gas, permitting the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the substrate.
During normal service of a gas turbine blade or vane, the airfoil layer is typically damaged by particle impacts and by oxidation/corrosion in the hot combustion gas environment. If the damage is not too severe, the gas turbine blade or vane may be removed from service, repaired, and returned to service. The repair typically includes, among other things, stripping away the damaged protective coating and thermal barrier coating layer, if any, from the external airfoil surface, and applying new protective coatings.
For those cases where the airfoil has internal cooling passages, the removal of the external protective coating during repair operations reduces the remaining structural thickness of base metal that lies between the internal cooling passage and the external airfoil surface. The inventors have determined that this is particularly a concern when the external protective coating, or an inner portion of the external protective coating, experiences significant diffusion into the base metal either during manufacturing or during service. For example, a typical 0.002 inch thick diffusion aluminide coating includes a diffusion zone that is about 0.001 inch thick, and an "add-on" layer that is about 0.001 inch thick. The diffusion zone consumes a portion of the wall of the airfoil, reducing its effective thickness for supporting loads. When this thickness becomes so reduced that it can no longer support the required structural loads, the turbine blade or vane becomes unrepairable, and must be discarded even though otherwise it could be repaired and returned to service.
There is a need for an improved approach to the protection of gas turbine airfoils containing internal cooling passages, which permits their repeated repair without loss of structural integrity. The present invention fulfills this need, and further provides related advantages.